Three-wing, six tilt-propulsion unit, VTOL aircraft

ABSTRACT

A vertical takeoff and landing aircraft having a fuselage with, preferably, three wings and six synchronously tilt-able propulsion units, each one mounted above, below, or on each half of the aforementioned three wings. The propulsion units are oriented vertically for vertical flight and horizontally for forward flight. Each propulsion unit comprises a propeller having a plurality of blades, where the pitch angle associated with the distal end of each blade and the proximal end of each blade are independently adjustable. As such, each of the propellers can be adjusted to exhibit a first blade pitch angle distribution optimized for vertical flight and a second blade pitch angle distribution optimized for forward flight.

REFERENCE TO RELATED APPLICATIONS

This patent application is a Continuation-in-Part Application of U.S.application Ser. No. 12/900,790 filed on Oct. 8, 2010, which isNon-Provisional Application of U.S. Provisional Application No.61/278,612 filed on Oct. 9, 2009; the disclosure of both priorapplications are incorporated herein by reference in their entirety.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The invention relates to the field of VTOL aircraft, and in particular,to an aircraft which is capable of sustained vertical flight followingthe loss of thrust from a propulsion unit. This means the continued safeflight following the failure of not only an engine but a gearbox orpropeller/rotor.

2. Description of Related Art

The advantages of an aircraft taking off and landing without an airportare obvious. Fixed wing aircraft are restricted to operations fromrelatively few airports and many of these airports are not the intendedfinal destination. In military applications, airbases are vulnerable toattack. The response time of forward positioning of a VTOL supportvehicle is significantly reduced and the range requirements are alsoreduced. Flexible positioning is also valuable in air ambulanceapplications, where the total response and delivery time “the goldenhour” is critical to the survival of the patient.

The pursuit of an aircraft which can take-off and land vertically whichpossesses the speed, range, and payload capacity of an airplane hascontinued since the invention of the helicopter.

Helicopter VTOL

The reason this search continues is that the helicopter is full ofperformance limitations and safety problems. The performance is limiteddue to its limited forward speed due to retreating blade stall. It alsohas a limited range due to its inefficiency compared to fixed wingaircraft. The limited range is further reduced by the utilization oflight weight turbine engines, which do not reach any reasonable fuelefficiency until operating at high altitudes where helicopters do notnormally operate. The helicopter must consume much of its power simplykeeping itself in the air, and approximately 15% of the power isconsumed by the tail anti-torque rotor just to keep the helicopter fromspinning. The helicopter must also deal with high vibration levels.There are many safety related problems that engineers and pilots arefully aware of. These are, loss of tail rotor effectiveness (LTE),Ground Resonance, Mast bumping, Loss of control during negative Gflight, Settling with power (VRS), Dynamic roll-over, anti-torque rotorfailures, and in case of engine failure, auto-rotations requiringunusually quick response by the pilot to maintain rotor speed and theimmediate requirement to select a landing site. So, the helicopter hasbeen the mainstay solution to our vertical flight needs because we havenot found a better solution, but the aviation industry has not stoppedsearching for better solutions.

Aircraft designers have been working on the concept of vertical takeoffand landing (VTOL) aircraft for many years. There are a number ofdifferent ways to combine vertical flight with the horizontal flight ofa conventional airplane, but developing a practical, redundant, andfixed wing performance hybrid aircraft has proven to be a surprisinglydifficult and elusive task.

The engineering challenge consists of achieving two main goals. Thefirst is to accomplish redundant and controllable vertical flight insuch a way that the very same mechanisms and equipment are required forforward flight. Any weight of exclusively vertical flight mechanisms isuseless during forward flight and represents a reduction in availablepayload relative to a fixed wing aircraft capability. The second goalconsists of achieving “power matching”. This simply means a successfulVTOL design should require the same power in vertical flight as forwardflight. Any mismatch represents excess capacity which corresponds toexcess weight in one mode of flight.

The present invention simultaneously achieves both goals.

Numerous approaches to VTOL aircraft have been explored over the years,of which the two Joint Strike Fighter concept demonstrators are only thelatest. There have been tilt-rotors like the V-22 Osprey, tilt-props andtilt-wings, as well as deflected-slipstreams, deflected-thrust, thrustaugmenters, and tail sitters. Many of these exotic designs have beentested at Edwards, and Air Force Flight Test Center pilots have oftenbeen called to fly some extremely unusual aircraft. In fact, there areso many attempted VTOL aircraft solutions to this problem that theresimply is not enough room in this patent application to list them alland their related problems. Much can be learned by studying the historyof these and other VTOL aircraft.

Bell Helicopter, Ryan, Hiller, GE, Lockheed, Curtis-Wright, LTV, andothers, mainly funded by the U.S. government have developed and testedvarious VTOL prototypes, known as X (for experimental) planes for years.I have listed just a sampling of the more promising designs. Manycrashed or failed continued support, a couple are with us today. Mostwere not able to achieve the two main VTOL goals explained earlier. Manyare in the literature and our aerospace museums.

Ducted Prop/Fan VTOL

Configurations known as Ducted Fans, both two and four duct versions,have been built and tested. The Navy X-22 (Bell Model D2127), with thefirst flight Mar. 17^(th) 1966, had four 1,267 HP engines with 7 footdiameter propellers mounted within four large ducts placed at each endof a forward and rearward wing. This craft first flew in 1966 andcrashed, on Aug. the 8^(th) 1966 due to a single propeller controlfailure. The last flight was 1988. The promise of higher static thrustefficiencies offered by ducted propellers or fans was offset by theadditional weight and drag of the large ducts during cruise flight. Thiscraft also had cross-linked drive-shafts to all four propulsion units toprovide redundancy in the event of engine failure. The weight of thiscomplex system and the ducts consumed much of the payload capability.The second prototype aircraft is currently on display in the NiagaraAerospace Museum, New York. The problem of a failure of a gearbox,shaft, or propeller and VTOL operation with airplane speed, range, andpayload was not solved by the ducted propeller or fan design.

The Army/Douglas/Doak VZ-4DA Tilt Ducted Prop research vehicle placed aducted fan at each end of a single conventional wing. Again, the ductgave the added static thrust efficiency for vertical flight but theextra weight and additional drag was a problem. The design used crosscoupled drive shafts which significantly increased weight. The problemof a failure of a gearbox, shaft, or propeller and VTOL operation withairplane speed, range, and payload was not solved by the ductedpropeller or fan design.

Lift Fan VTOL

On Aug. 13, 1965, Maj. Robert L. Baldwin lifted an oddly humpbackedbrown jet into the air and began the Air Force Flight Evaluation of theGE-Ryan XV-5A. General Electric had been researching a fan-in-wingconcept for V/STOL aircraft, and late in 1961 it won an Army contractfor a concept demonstrator. GE subcontracted the design and constructionwork to Ryan. The XV-5A that resulted was a small, fighter-like design:44 feet long with a 30-foot wingspan.

A pair of J85-GE-5 turbojets mounted within the fuselage providedapproximately 5,000 pounds of thrust in normal flight. When verticalthrust was needed, the pilot could actuate a diverter valve thatdirected some of the exhaust gases to a pair of fans, 5 feet indiameter, located in the inboard portion of each wing. The wing fansrotated in opposite directions and were covered by large hinged doors inconventional flight. Exhaust gas also powered a smaller fan in the nosethat provided pitch control and a measure of additional lift. All threefans together provided 16,000 pounds of vertical thrust. A set oflouvered vanes underneath each of the large wing fans could vector thethrust in any direction and provided yaw control. Much of the allimportant payload capability was consumed by the complex fan system. Thetransitional behavior was reported as quite “abrupt” preventing it frombeing adequate for “service pilots”. The XV-5A crashed at Edwards AirForce Base killing test pilot Lou Everett. The problem of a failure of afan and VTOL operation with airplane speed, payload, and range was notsolved by this ducted fan design.

Tilt Wing VTOL

Configurations where the whole wing tilts (Tilt Wings) with engines andpropellers attached have been built and tested. On Jan. 27, 1961 theNavy Bureau of Naval Weapons (Bu Weps) created the Tri-Service AssaultTransport Program. The original outline had been drawn up as areplacement for the Sikorsky HR2S, with a payload of 10,000 pounds,operational radius of 250 miles, and cruising speed of 250-300 knots.This is roughly what the Marine Corp has needed for many years and hasculminated in the requirements for the V22 Osprey twin Tilt Rotoraircraft. The Ling-Temco-Vought (LTV) of Grand Prairie Tex. XC-142(A)cargo airplane (actually known as the Vought-Ryan-Hiller XC-142, butwhen Vought became part of the LTV conglomerate the name was dropped) isthe most notable tilt wing VTOL. The XC-142A first flew on Sep. 29,1964, and on Jan. 11, 1965, it completed its first transitional flight.My father was a design engineer at LTV when I watched this plane as ateenager. I was there when it crashed in the marshes of the lake nearour home killing three test pilots due to a tail rotor drive shaftfailure. This aircraft had problems with cross-linked drive shaftvibrations. It also had problems with wing angles greater than about 35degrees. The wake due to wing stall caused serious control problems andas it converted from forward flight, the increase in wing angle ofattack yielded too much lift and the aircraft climbed. These behaviorsand problems were judged unacceptable. This aircraft configuration didnot solve the problems resulting from the failure of a gearbox,driveshaft or propeller.

The Vertol V-76 is an additional example of a Tilt Wing researchaircraft. This aircraft configuration did not solve the failure of agearbox, driveshaft or propeller problems.

Deflected Thrust VTOL

Bell constructed the X-14 as an open-cockpit, all-aluminum monoplane. Itwas powered by two Armstrong Siddeley Viper turbojet engines equippedwith thrust deflectors sited at the aircraft's center of gravity. Theengines are fixed in position; transition from vertical to horizontalflight is achieved with a system of movable vanes that control thedirection of engine thrust. Top speed was 180 miles per hour with aservice ceiling of 20,000 feet. The X-14 was designed using existingparts from two Beechcraft aircraft: wings, ailerons, and landing gear ofa Beech Bonanza and the tail cone and empennage of a Beech T-34 (amilitary trainer variant of the Bonanza).

The X-14 first flew on 19 Feb. 1957 as a vertical takeoff, hover, thenvertical landing. The first transition from hover to horizontal flightoccurred on 24 May, 1958. In 1959, its Viper engines were replaced withGeneral Electric J85 engines. That year the aircraft was also deliveredto the NASA Ames Research Center as the X-14A. It served as a testaircraft with NASA until 1981. The X-14 project provided a great deal ofdata on Vertical Take Off and Landing aircraft. The X-14A was asuccessful research aircraft. The VTOL design did not produce apractical aircraft with meaningful payload and range.

Directed Jet Thrust VTOL

Configurations using the Directed Thrust of jet engines have beentested. A small single seat Bell VTOL experimental had small turbojetengines placed on each side of the fuselage underneath a high wing. Theengines were tilt-able for conversion to forward flight. Reactionnozzles, using bleed air provided directional control during hovering.The problem of engine failure was not solved by this aircraft design.

The MBB VJ-101 Jet Lift vehicle consisted of an F104 aircraft withtilt-able turbojet engines placed at each end of its conventional wing.The problem of engine failure was not solved by this aircraft design.

The British developed the fore-runner to the Marine's AV-8 Harrier,which proved itself in the Falkland's Islands invasion. But, to be trulyuseful, this aircraft must make a short conventional takeoff to carryits all important payload, making this vehicle a STOVL, Short Takeoffand Vertical Landing Craft. True single engine VTOL solutions usingdirected jet thrust will seldom be efficient or inexpensive, certainlynot safe.

The AV-8 Harrier uses an ejection seat as a solution for engineproblems. This is a very narrow solution initially designed for forwardair support of ground troops.

The problem of a failure of a propulsion unit is not solved by thesedesigns. The tilt jet embodiment of the present invention solves theproblems of previous tilt jet configurations.

Jet Thrust Augmentation

Lockheed set up a privately funded VTOL design that was approached in adifferent way, using augmented lift-jets (thrust augmentation). It mixedhot engine exhaust gas with cold air to increase engine thrust. Lockheedmade and tested their Model 330 Hummingbird where the exhaust from thejet engine was directed through ducts or nozzles that were installedwithin the fuselage. Most of the fuselage space was occupied as themixing chamber with cold air intake from the upper hatch. Lockheedproposed this study to U.S. Army and was awarded a contract for twoprototypes, designated XV-4A in competition with Ryan XV-5A and HawkerXV-6A in 1961. The first conventional flight was on July 1962. Threemonths later, first hover was tested but the actual thrust augmentationwas much less than expected. The first XV-4A crashed in 1964. The secondXV-4A then was modified as XV-4B for flight research by USAF until itcrashed in 1969. The problem of engine failure and useful payload wasnot solved by this aircraft design.

Tail Sitter VTOL

In this configuration the aircraft literally sits vertically on itstail. This is accomplished by having three or four tail surfaces withsmall wheels attached at their ends. The aircraft normally has a counterrotating propeller located on the nose. The pilot must climb up aplatform to enter the cockpit. These versions were built and flown butall test pilots had serious visibility problems when attemptinglandings. They were intended for maritime service aboard ships withheaving decks! They actually accomplished the goal of vertical andhorizontal flight using the same equipment and power but were notcapable of carrying any real payload other than the pilot and fuel.Their propeller disk loadings were simply too great for any meaningfulpayload and the landing challenge was completely unacceptable for anymanned service.

Tilt Rotor/Prop VTOL

Designs using Tilted Propellers (Tilt-Props) are not new, theTri-Service/Curtiss-Wright X-19 Tandem Wing Radial Force Tilt Propaircraft placed Tilt Props at each end of two wings, creating a two wingfour Tilt-Prop design. This aircraft used a cross-coupled drive shaftsystem to provide for an engine failure. The problem of a failure of agearbox, shaft, or propeller is not solved by this design.

In October of 1953, Bell Helicopter was awarded a development contractto produce two aircraft for testing purposes. The Bell Helicopter XV-3(originally known as the XH-33, classifying it as a helicopter) (Bellmodel 200) (the designation was changed in 1962 to XV-3A) first flewAug. 11, 1955, achieved full conversion in 1958, was developed toexplore the convertiplane concept and eventually successfullydemonstrated the Tilt Rotor concept. These aircraft are referred to asTilt Rotors, referring to the propellers as rotors, because they aremainly helicopter rotors with similar control heads. This vehicleconsisted of two tiltable helicopter like rotors placed at each end of awing, driven by a single engine. The prototype encountered significantvibrations during the transition to forward flight. These vibrationswere eventually reduced. The data and experience from the XV-3 programwere key elements used to successfully develop the Bell XV-15. Theproblem of a failure of a gearbox, shaft, or propeller/rotor is notsolved by this design.

The development of the XV-15 Tilt-rotor research aircraft was initiatedin 1973 with joint Army/NASA funding as a “proof of concept”, or“technology Demonstrator” program, with two aircraft being built by BellHelicopter Textron (BHT) in 1977. Ship number 1 was given NASA number702, and ship #2 was 703. Aircraft development, airworthiness testing,and the basic “proof of concept” testing were completed in September1979.

The aircraft is powered by twin Lycoming T-53 turbo-shaft engines thatare connected by a cross-shaft and drive three-bladed, 25 ft diametermetal rotors (the size extensively tested in a wind tunnel). The enginesand main transmissions are located in wingtip nacelles to minimize theoperational loads on the cross-shaft system and, with the rotors, tiltas a single unit. For takeoff, the prop-rotors and their engines areused in the straight-up position where the thrust is directed downward.The XV-15 then climbs vertically into the air like a helicopter. Theproblem of a failure of a gearbox, shaft, or rotor is not solved by thisdesign.

The BA609, Bell-Agusta Tilt Rotor is the commercial offshoot of the BellXV-15, the tilt rotor proof of concept vehicle built at the main Bellplant in Hurst Tex. and tested at the Arlington Tex. facility. The BA609is still in development as of this writing and is not expected untilabout 2012. However, like the XV-15, this aircraft uses a cross-coupleddrive shaft system in case of engine failure. Thus, the problem of afailure of a gearbox, shaft, or rotor is not solved by this design.

The Bell/Boeing V22 Osprey Tilt Rotor is the military offshoot of theBell XV-15 tilt-rotor. It uses two 6,150 HP engines and cruises at about241 kts. It was intended to lift 25 fully loaded combat troops andquickly carry them 500 miles to fight its way in and out of high threatlanding zones. This vehicle represents a solution to Tilt Rotor flight,but has been quite costly in terms of development and unit cost. Theloss of a V22 in Arizona during a simulated mission is suspected to bedue to asymmetrical Vortex Ring State (VRS). This is when one rotoreffectively loses lift and the aircraft rolls and plunges to the ground.The V22 utilizes cross coupled shafts, so that the failure of one engineallows the remaining engine to power the opposite rotor. This iscomplex, heavy and expensive. This may provide a solution for a singleengine failure, but not a gearbox, bearing, rotor, or blade failure.This may be a reasonable risk to the military, compared to the lives itcan save, but may prove problematic and too costly in the commercialmarket.

Twin Tilt Rotor Aircraft Problems

The current technology twin tilt rotor V22 Osprey's recent performanceis used as the best example of twin tilt rotor problems. Still,examination of the V22 performance data makes it clear that this is nota VTOL aircraft, but instead a STOVL aircraft.

1. The failure of either propulsion unit to produce thrust, results inthe loss of the aircraft and occupants. The cross coupled drive shaftsystem provides a marginal performance backup limited to engine failure,not gearbox or rotor failures. Moreover, the installed weight of thecross-coupled drive shaft system reduces the all-important payloadweight. The loss of 50% of this aircraft's power results in its'inability to continue its' mission and in many circumstances requiresimmediate landing. The present invention solves this problem with its'fundamental distributed propulsion design, providing for continuedflight following the complete loss of thrust by any propulsion unit, nomatter what the cause.

2. Inability to use evasive maneuvering tactics to avoid hostile fire,similar to that found in high threat landing zones. Quick maneuveringhas resulted in the cracking and breaking of rotor root components whichmay result in the loss of a blade, and therefore the aircraft. The 38foot rotors are limited to about 10 degree flapping angles. Quickmaneuvering can (and has) easily exceed this angle resulting in thefailure of a rotor and the potential loss of the aircraft. This limitedmaneuverability might be acceptable for commercial operations, but notfor the intended applications of the military. The V22 is intended forinsertion and extraction missions during ground fire. The presentinvention solves this problem by using much smaller diameter rigidpropellers as opposed to helicopter like rotors which allow the bladesto flap and distribute the load to six propellers instead of two rotors.The propellers are smaller for the same disk loading, therefore theyhave a smaller moment of inertia. Propellers have been used successfullyon aerobatic airplanes and WWII dog fighters without blade separationdue to these high gyroscopic forces. For example, the V22 Osprey has tworotors totaling a disk area of 2,268 square feet. Each blade is roughly19 feet long (including hub radius), the length of a blade in a sixpropeller system of the same area would be eleven feet. This representsa significant reduction in blade root forces due to high G maneuvers.

3. Limited center of gravity (c.g.) range. The V22 design is inherentlya narrow C.G. aircraft. The side by side location of the rotors doesnothing to improve longitudinal c.g. range and the V22 has the highlylimited c.g. range of a helicopter with a single rotor. It was intendedthat soldiers “down rope” from the left and right side of the V22, butthe rotor wash blew them off the ropes due to the hurricane like windbelow the rotors. A single rope was attached below the tail above therear ramp area avoiding the rotor wash. This limited C.G. range preventssoldiers from moving to the rear ramp during “down roping” insertions,until the previous two have released the ropes' end at the ground. Thissignificantly increases the time required to down rope 25 marinesexposing the V22 to hostile ground fire for an unusually long period. Ifsoldiers are excited and rush to the rear of this craft exceeding itsaft C.G. limit, control will be lost along with the aircraft. Thepresent invention solves this problem by distributing the thrust aroundthe aircraft. The C.G. range is quite large and is therefore verytolerant of moving payload. In addition, the present motivation canaccommodate simultaneous right and left down roping because the middlewing shields the soldiers from most propulsion unit down washoriginating from the wing ends.

4. Settling with power or Vortex Ring State (VRS) is a well known (byengineers and pilots) problem with helicopters. It is a condition wherean annulus or “donut” of accelerated air is created around the actuator(rotor) disk. Simply put, it's what happens when you try and fly in yourown rotor wake. Roughly speaking, the air is turbulent and you begin tolose lift, if you raise the collective (increasing pitch) to correct thesituation, it becomes worse. It is similar to running up the downescalator. The solution has been to reduce collective and push thecyclic forward to fly out of the disturbed air. An accepted method toavoid it is, as an example, is not to descend more than 300 fpm if youare flying less than 30 kts. When a helicopter gets in this situation,it can recover as long as it has enough altitude. Even tandem rotorhelicopters can experience this behavior. The front rotor normallyenters VRS first causing the front to drop, assisting in the necessaryforward motion for recovery. The V22 configuration is different. Whenone of its rotors enters VRS, the aircraft begins to roll, the pilotinstinctively corrects with opposite stick, worsening the problem andthe ship rolls over and plunges toward the ground. This is what possiblyhappened near Tucson, Ariz. on a simulated night mission killing manyMarines. The present invention solves the VRS susceptibility problem forTilt Propeller configurations by distributing the load among sixpropellers. When a propeller enters VRS the present invention shouldremain stable and controllable.

SUMMARY OF THE INVENTION

In accordance with one aspect of the present invention, there isdisclosed a vertical takeoff and landing aircraft comprising a fuselage,a plurality of wings attached to the fuselage, a first rotatablepropulsion unit attached to a first side of each of the plurality ofwings and a second rotatable propulsion unit attached to a second sideof each of the plurality of wings. Each rotatable propulsion unitcomprises a propeller and a propeller hub. Moreover, each propellercomprises a plurality of blades, where each of the plurality of bladeshas a hole along its longitudinal axis. Each propeller also comprise aplurality of rotatable rods, where each of the plurality of rods extendsinto the hole of a corresponding one of the plurality of blades. Theproximal end of each of the plurality of blades is enclosed in oradjacent to the propeller hub and rotatable around the corresponding oneof the plurality of rotatable rods. The distal end of each of theplurality of rods is fixed to a distal end of the corresponding one ofthe plurality of blades.

Before explaining at least one embodiment of the invention in detail, itis to be understood that the invention is not limited in its applicationto the details of construction and the arrangements of the componentsset forth in the following description or illustrated in the drawings.The invention is capable of other embodiments and of being practiced andcarried out in various ways. Also it is to be understood that thephraseology and terminology employed herein are for the purpose ofdescription and should not be regarded as limiting.

As such, those skilled in the art will appreciate that the conception,upon which this disclosure is based, may readily be utilized as a basisfor the designing of other structures, methods and systems for carryingout the several purposes of the present invention. It is important,therefore, that the claims be regarded as including such equivalentconstructions insofar as they do not depart from the spirit and scope ofthe present invention.

The foregoing has outlined, rather broadly, the preferred features ofthe present invention so that those skilled in the art may betterunderstand the detailed description of the invention that follows.Additional features of the invention will be described hereinafter thatform the subject of the claims of the invention. Those skilled in theart should appreciate that they can readily use the disclosed conceptionand specific embodiment as a basis for designing or modifying otherstructures for carrying out the same purposes of the present inventionand that such other structures do not depart from the spirit and scopeof the invention in its broadest form.

BRIEF DESCRIPTION OF THE DRAWINGS

Other aspects, features, and advantages of the present invention willbecome more fully apparent from the following detailed description, theappended claim, and the accompanying drawings in which similar elementsare given similar reference numerals.

FIGS. 1 a and 1 b are top views of the aircraft with propulsion units inthe horizontal and vertical positions;

FIGS. 2 a and 2 b are side views of the aircraft with propulsion unitsin the horizontal and vertical positions;

FIGS. 3 a and 3 b are front views of the aircraft with propulsion unitsin the horizontal and vertical positions;

FIG. 4 is a Flight Control Thrust Vectoring Mechanism; and

FIG. 5 is a cross-sectional view along the longitudinal axis of apropeller 80, in accordance with an exemplary embodiment of the presentinvention.

DESCRIPTION OF THE PREFERRED EMBODIMENT

The invention is a vertical takeoff and landing aircraft. When thepropulsion units consist of engine driven propellers, this VTOLconfiguration is known as a tilt-prop aircraft, since the propellers aretilted forward for forward flight and tilted vertically for verticalflight. When the propulsion units consist of engine driven rotors, thisVTOL configuration is known as a tilt-rotor aircraft. When thepropulsion units consist of engine driven fans, this VTOL configurationis known as a tilt-fan aircraft.

The term “Propulsion Unit” as used within this document refers to anymethod of producing thrust. The example of an engine driven propeller ischosen solely for illustration purposes and is not intended to limit thescope of this invention. The invention is valid for alternate means ofpropulsion including jet engines. When the propulsion units consist ofjet engines this VTOL configuration is known as a “Tilt Jet” since thejets are tilted forward for forward flight and tilted vertically forvertical flight. The invention is valid for alternate propulsion unittilting implementations. The engine (piston, turbine, rotary, electricor other power source) may be mounted on the wing and transfer powerthrough a gearbox into its tilt-able propeller or rotor. This designallows the engine or motor to remain relatively fixed in a singleposition without having to operate in multiple positions of thepropeller.

The aircraft configuration consists of a conventional aircraft fuselage,with a nose, with or without pilot and/or co-pilot crew stations in thecase of Unmanned Aerial Vehicle (UAV) applications, a central cabin orpayload area, and a tapering empennage. The aircraft has three wings,the front wing, middle wing, and the rear wing. Two propulsion units aremounted above, below, or on each of the three wings, yielding sixpropulsion units. The wings are fixed to the fuselage and the propulsionunits rotate in unison to either of two (not including intermediate)positions, vertical or horizontal.

For electric motor, rotary, piston or turbine driven propeller or rotorembodiments, the propulsion units on opposite sides of the aircraft turnin opposite directions to cancel rotational moments about the yaw axisdue to propeller or rotor torque. Small flapped wing panels are fixedoutboard of the forward and rearward propulsion units. These wing panelsare located within the propulsion units' propeller slipstream. Theyprovide yaw control during vertical flight. Their flaps are disabled inthe neutral position once the propulsion units advance toward thehorizontal position.

The main landing wheels are located at the rear end of the forwardpropulsion units. A retractable and steerable tail wheel is located onthe center line of the fuselage near the rear of the aircraft andretracts rearward and upwards into the normally unused space in the tailcone or alternatively for applications which require rear doors or aramp the tail wheel may retract forward into the bottom of the fuselage.This eliminates the normally complex and heavy landing gear retractionand extension system increasing payload capacity. When the engines ormotors rotate with the propellers or rotors the main wheels are mountedto the aft end of the forward propulsion unit engine support structures.This takes advantage of the existing structural load path which alreadyexists for the engine support. When the engines or motors are mounted ina fixed position with the propellers or rotors tilting, the main wheelsmay be attached to the aft end of the tilting assembly. The propulsionunits are spaced further apart than typical main gear designs increasingground stability. Gear up landings are not possible with this inventionas the landing gear is always down when the aircraft is in verticalflight mode. Separate landing gear controls and systems are notrequired. Proper placement of the main gear below the nacelle centerline and clam shell gear doors can enable partial conventional takeoffand landings (CTOL) to enable additional payload capability when arunway is available. This is accomplished by placing the propulsionunits in an intermediate position considering ground clearance isprovided for the propeller, fan, or rotor tips. With jet propulsionunits this would not be a problem.

For jet engine propulsion embodiments, yaw control is accomplished byexhaust deflection or bleed air supplied attitude control nozzles,methods instead of the yaw control wing panel required in the tiltpropeller embodiment. The main landing gear consists of the same systemdescribed above except that the main wheels are not placed within thejet exhaust at the rear of the propulsion units but placed below theexhaust area.

The flapped front wing, sometimes called a canard, and rear flapped wingoperate differentially providing pitch control. The aircraft contains aconventional vertical stabilizer and rudder assembly. The rudderprovides conventional yaw control during forward flight. The middle wingcontains conventional ailerons for roll control. When manned, the crewstation(s) contain conventional helicopter controls, namely, acollective control used in vertical flight mode, a cyclic control forpitch and roll control, and rudder pedals for yaw control.

The forward wing is set at an effective angle of attack greater than themain and rear wing. This assures that this forward wing stalls first,dumping its load and causing a nose down pitching moment, for safety,restoring proper flight attitude, reducing the chance of the middle andrear wings stalling. Additional advantages of providing pitch control atthe front and rear wing, as opposed to the single conventional rearelevator is the elimination of trim drag, the normal elevator down forceand total elevator loss of control which can occur in deep stalls. Someaircraft have airfoils known as strakes, which are not required in thisembodiment, mounted below their tails to provide a nose down pitchingmoment to prevent this from happening.

The rear wing may be mounted above the middle wing and the middle wingabove the front wing. This arrangement reduces the exposure of each wingfrom flying in the downwash of the wing ahead of it, decreasing drag.The wing spans of each wing are chosen to provide the design wing spanand position each set of propulsion units such that the thrust wake offorward units do not disturb the propulsion units that are mounted totheir rear.

While in vertical flight, the six propulsion units are arranged aroundthe aircraft producing thrust. Imagine a round table with six legs.Remove one leg and the table remains standing! The center of gravity ofthis aircraft is generally located about its' center. The propulsionunits are placed such that the remaining thrust following the loss ofthrust from one propulsion unit will maintain longitudinal and lateralstatic stability, therefore supporting the aircraft.

So, the weight, cost, and non-reliability of cross coupled engine driveshaft systems required in previous discussed designs is not required forcontinued flight following an engine failure in this invention. Eachpropulsion unit contributes only ⅙ of the total thrust. The twin tiltrotor V22 Osprey and BA609 engines must be sized such that one enginemust supply the total power required for vertical flight. This meansthese engines need to be capable of greater than 200% of the normallyrequired power. The additional reserve power required of these enginesrepresents a lot of extra pounds and dollars. They must carry this extraweight of the engines and drive shafts all the time which reduces theirpayload capacity. The owner or operator must pay the initial cost, andthe continuing maintenance, and overhaul costs associated with thisexcess capacity. Also notice that the failure of a gearbox, rotor systemor blade and asymmetrical VRS will result in the total loss of theseaircraft.

The present invention requires its' propulsion units to have reservepower, to replace the thrust from the failed propulsion unit, but farless than previously discussed configurations. Much of this reservepower is already required for normal de-rating for reliability, aging,additional power for control, and additional power for vertical climbingand vertical decelerations. So, this distributed power inventionrequires little additional capacity due to a failed propulsion unit.

This invention solves the critical problem of failed engines, gearboxes,propellers, or rotors resulting in the loss of the aircraft andoccupants.

This invention solves the problem of limited maneuverability of twintilt rotor designs.

This invention solves the problem of limited center of gravity range ofsingle or twin tilt prop/rotor designs.

This invention overcomes the susceptibility to Vortex Ring State ofhelicopters and twin tilt prop/rotor designs.

This invention solves the problem of propulsion redundancy requiringextra capacity and weight.

This invention solves the problem of a propulsion unit failurenecessitating the eminent requirement to land.

This invention solves the problem of truly redundant VTOL flightcombined with the speed, payload, and range similar to fixed wingaircraft.

This invention significantly reduces the retractable tricycle landinggear installed weight and complexity resulting in significantlyincreased payload capacity.

Shown in FIGS. 1-4 are a top view, a side view, a front view of theaircraft and a Flight Control Thrust Vectoring Mechanism for a VTOLtilt-propulsion aircraft in accordance with the principles of theinvention.

Large diameter tilt rotor propulsion units are best for relatively heavylift and lower speed applications similar to the V22 Osprey due to theirintrinsic lower disk loading, yielding higher lift efficiency and higherrotor drag during cruise. The large blade areas of these rotors,requires large power to drag them through the air at high speeds.Although these aircraft normally contain excess power allowing them toachieve high speed flight, the fuel consumed at these speedssignificantly decreases range. The V22 Osprey is capable of flight at300 kts, but recommended cruise is 241 kts.

Tilt propeller propulsion units are best for medium lift applicationsrequiring relatively higher cruise speeds, similar to turboprop fixedwing aircraft, due to the propellers' lower weight and drag at highercruise speeds. In contrast, tilt jet propulsion units are best forhighest speed applications.

The choice between reciprocating engines and turbo-shaft engines, as thepower source, is not so simple. Where cost and propulsive efficiency,for range, are important, reciprocating engines will be the bestcandidate. Where higher power, speed, and lighter weight become adriving factor, turbo-shaft engines will be the best candidate. Theslightly lower reliability of modern reciprocating engines, relative toturbines, is no longer a factor in the present invention due to itsredundancy. Small UAVs will use reciprocating engines. Other UAVs mayuse electric motors. When the UAV must carry ordinance or significantpayload such as sensor weights exceeding 500 pounds, turbo-shaft enginesmay be indicated. Commercial market air ambulances will use turbo-shaftengines. The main variations of the invention involve this choice ofengine size and propulsion technology. The choice is determined by asimple trade study considering the intended application and theavailable engines. As long as the embodiment of the present inventionprovides for the static stability of the aircraft following the loss ofthe thrust of a single propulsion unit, and the thrust wake from eachpropulsion unit does not materially affect the performance of propulsionunits to its rear, and the center of gravity range of the aircraftdesign remains within a polygon described by the interconnection oflines between thrust producing propulsion units, then the inventions'requirements are materially met.

Referring to FIGS. 1-3, there is shown an aircraft 10 with sixpropulsion units 26 which can be in the vertical position for take offand landing and in the horizontal position for forward flight inaccordance with the principles of the invention. The tilt-prop aircrafthere disclosed, while not limited to any specific application, isintended to satisfy the medium weight and medium speed requirement.

The aircraft has three wings. The rear wing 12 is mounted above themiddle wing 14 which is mounted above the front wing 16. The front wing16 and rear wing 12 have differentially connected flaps 18 which providethe conventional forward flight pitch control. The front wing 16 is setat a higher effective angle of attack which assures stall prior to themiddle wing 14 and rear wing 12.

The middle wing 14 contains conventional ailerons 20 for roll control inforward flight. The middle wing may extend beyond the propulsion unitswhen a higher aspect ratio wing is required.

The aircraft has a conventional vertical stabilizer 22 and rudderassembly 24.

Located on each wing half is a propulsion unit 26 where the propulsionunits of the aircraft of the FIGS. 1 a through 3 a are horizontallyoriented and the propulsion units of the aircraft of FIGS. 1 b through 3b are vertically oriented.

Located outboard of the front and rear propulsion units 26 are yawcontrol panels 32 with flaps 30. They are fixed to the propulsion unitsand rotate with them providing yaw control during vertical flight.

The rear wing 12 with propulsion units 26 may be mounted at the top ofthe vertical stabilizer 22 as a traditional T-tail arrangement, mountedin the middle as a traditional cruciform arrangement, or at the bottom.Mounting the rear wing 12 at the bottom of the vertical stabilizer 22may require a jog in the empennage, similar to the V22 empennage, tomaintain the rear wing 12 mounting position above the middle wing 14.Alternatively, the middle wing 14 may be located in the middle of thefuselage 28, allowing the rear wing 12 to be mounted on the uppersurface at the base of the vertical stabilizer 22. This places themiddle wing 14 spar in the cabin and may require an undesirable heavyring carry through structure to prevent this. A UAV may use this methodwithout the ring structure. Asymmetric thrust may be used for forwardflight yaw control as long as the vertical stabilizer is sized forstatic lateral axis stability.

Analysis by the applicant of the invention has produced these listeddesign goals:

1. The engines' power should be approximately 150% of the power requiredfor the “Hover Out of Ground Effect” H.O.G.E. hover and temperaturerequirements. This will provide for reliability de-rating, reduction ofthrust due to engine aging, requirements for control power, and heavepower (vertical acceleration and deceleration). This reserve generallyshould power match that necessary for engine-out operations.

2. All unnecessary weight should be avoided. The weight growth factorfor VTOL designs is quite large.

3. Lightweight engines should be specified. Traditional horizontallyopposed piston Lycoming or Continental engines may not be acceptable.Their horsepower to weight ratios of 1 to 2 are not appropriate for VTOLTilt-Prop applications.

4. Complex and weighty propeller systems are not appropriate. Selectlight weight composite designs.

5. Normal utility load factors are acceptable.

6. Retractable landing gear, either tired tricycle or retractablelanding skids are indicted for the highest speed requirements. Wellfaired fixed gear is appropriate up to about 200 kts. The gear should bedesigned for taxi capability and VTOL operations only. The disclosedmethod of placing the main wheels at the aft end of the forwardpropulsion units will provide the lowest cost and greatest payload.

7. Propeller power factors greater than 6 pounds thrust per horsepowershould be sought.

8. Propeller disk loading of about 15 pounds per square foot or lessshould be sought.

9. Consider light weight propellers with novel blade twist technologies.These offer blade angle distributions favorable to both hover and cruiseefficiency requirements.

10. Blade tip speeds during hover should be about 800 fps. Tip speedsduring cruise should not exceed 0.8 Mach, except in tilt fanapplications.

11. The propeller, gearbox, and engine combinations' static performanceshould be fully characterized and tested before the airframe detaildesign begins.

12. Power matching the cruise and hover requirement should be a designgoal. When necessary, consider reducing the cruise speed requirement tomeet this goal.

13. The rear wing should be mounted above the middle wing which shouldbe mounted above the front wing to reduce drag due to downwash.

14. The front and rear wing should be flapped and differentiallycontrolled to provide the conventional forward flight pitch control, andthe front wing should be set at an effective angle of attack whichassures stall prior to the middle and rear wings.

15. The middle wing should contain conventional ailerons for rollcontrol in forward flight.

16. The aircraft should contain a conventional vertical stabilizer andrudder assembly. The rear wing with propulsion units may be mounted atthe top as a traditional T-tail arrangement, mounted in the middle as atraditional cruciform arrangement, or at the bottom. Mounting the rearwing at the bottom of the vertical stabilizer may require a jog in theempennage, similar to the V22 empennage, to maintain the rear wingmounting position above the main middle wing. Alternatively, the mainmiddle wing may be located in the middle of the fuselage, allowing therear wing to be mounted on the upper surface at the base of the verticalstabilizer. This places the middle wing spar in the cabin or requires anundesirable heavy ring carry through structure. Asymmetric thrust may beused for yaw control as long as the vertical stabilizer is sized forstatic lateral axis stability.

17. The aircraft should have a highly reliable and simple thrustvectoring control mixer. The system should be analyzed carefully toeliminate single point failures causing total system failure. Failsafethrust positions at the engines are necessary. Consider propulsion unitfailsafe to some intermediate thrust by analyzing the hover powerrequired at minimum useful load (single pilot and low fuel). Considerincluding anti jam links in the mixers' thrust command outputs.

18. The aircraft design team should take advantage of the sixalternators and batteries located within the propulsion units, to obtaina fail proof electrical power source for the avionics systems. A minimumrequirement is a left and right main electrical buss with a cross tiecontactor. The busses should be located on opposite sides of theaircraft.

19. The fuel tanks should remain within the propulsion unit nacelles forsafety. When additional fuel is required, each wing can have fuelbladders. A single point fueling system is heavy and brings the fuelinto the fuselage. This should only be considered in UAV or militaryapplication where it might be mandatory. For commercial applications itis not abnormal for four or more tanks to require filling.

Control of the Aircraft During Forward Flight

The aircraft is controlled conventionally by helicopter type flightcontrols in forward flight. The collective flight control is not usedduring forward flight. Pitch and roll is controlled by the cyclic flightcontrol which is differentially connected to the front and rear wingelevators for pitch control and middle wing ailerons for roll control.The rudder pedals provide yaw control and are connected to the verticalstabilizer rudder.

Control of the Aircraft in Vertical Flight

The control of motion about the pitch and roll axis is by way of thrustvectoring, which is accomplished by reducing the propulsion units'thrust in the direction you want the vehicle to move toward andincreasing the propulsion units' thrust opposite this direction. Thismay be accomplished by a mechanical, electronic analog, digital orhybrid method, however, the following example is mechanical.

This is illustrated by controlling engine RPM, by way of throttling.Thrust may be controlled by differentially controlling propeller pitch,engine RPM or many other methods.

Referring to FIG. 4, there is shown a mechanism 50 which converts cyclicstick movement to thrust commands to six propulsion units. Thismechanism is used to control the propulsion unit's thrust resulting inthe pitch and roll of the vehicle in vertical flight. The mechanism isfundamentally a mechanical implementation of a rectangular (X, and Y) topolar (Displacement, and Angle) (Rho, Theta) coordinate converter, withthe exception of an additional input, Z 68. This Z input 68 causes anequal change in six displacement outputs 60. Therefore, there are threelinear inputs, X 76, Y 64, and Z 68 and six displacement outputs 60. Letthe X input 76 be the pitch (fore or aft) position of the cycliccontrol, the Y input 64 be the roll (right or left) position of thecyclic control, and the Z input 68 be the position of the collective (upor down) control.

The mechanism 50 consists of a central two piece vertical rod assembly,consisting of two coaxially oriented rod pieces 52, 54, one above theother, with a universal joint 56 connected between them. The bottomvertical rod segment 54 is mounted in a fixed position linear bearing 66near its bottom end, allowing vertical movement of the rod assemblywhich consists of 52, 54, 56, 58, 60, 62, 64, 68, 70, 74, and 76. A disk74 is fixed on rod 54 above linear bearing 66 and below universal joint56. Attached to the periphery of disk 74 is the Z axis input fitting 68.This is the Z axis, or collective input to the mechanism. When you raiseor lower the collective control the rod assembly rises and falls inunison. Disk 74 contains a vertically oriented fixed linear bearing 70with a small fixed vertical rod 72 passing through it. This prevents therod assembly 52, 54, 56, 58, 60, 62, 64, 68, 70, 74, and 76 fromrotating. A disk 58 is mounted on the upper rod 52 segment midwaybetween the universal joint 56 and top end. There are six attachmentpoints 60 around disk 58 which correspond to the relative angularlocations of the propulsion units on the aircraft. These six attachmentpoints 60 are the six thrust command outputs of the mechanism. A similardisk 62 is attached to the top of the upper rod segment. This disk hastwo orthogonally oriented attachment points 64 and 76 on the disk. Theseare the X 76 or pitch, and Y 64 or roll inputs. The cyclic control isconnected to these two inputs, X 76 or pitch, and Y 64 or roll, and thetop of the vertical rod 52 mimics the cyclic control position.

In operation, moving the cyclic flight control toward the right frontdirection, say 45 degrees to the right, causes the disk 62 at the top ofthe mechanism 50 to move forward and to the right since the disk 62 isconnected to the cyclic control by way of the Pitch “X” 76 and Roll “Y”64 inputs. The centrally located thrust command disk 58 mounted abovethe universal joint 56 moves with the rod 52 in the same direction, 45degrees to the right. The thrust command attachment points 60 also movebecause they are part of the disk. Assuming the use of propeller pitchcontrol cables, the thrust outputs 60 from the disk 58 push (reducingpitch and therefore thrust) on the cables in the direction of thedesired vehicle movement and pull (increasing pitch and thereforethrust) on thrust cables in the opposite direction in an amountproportional to their relative angular placements.

Raising the collective flight control causes the rod assembly 52, 54,56, 58, 60, 62, 64, 68, 70, 74, and 76 to rise and therefore pulls onall thrust cables equally, increasing the propeller pitch and thereforethrust on all six propulsion units equally.

Yaw control is provided through rudder pedals, controlling the flappedoutboard wing panels. The conventional rudder remains connected andmoves with the rudder pedals during hovering flight. Roll and pitchcontrol by thrust modulation remains active until the propulsion unitsare commanded to leave the vertical position. The front and rear wingpitch control flaps and middle wing ailerons remain connected to thecyclic flight control during hovering flight. The front, middle, andrear wing flaps may be placed in a down position to reduce the verticaldrag on the wing caused by the high velocity propeller wakes. Theseflaps must return to normal forward flight operation prior to disablingthrust vectoring control.

Control During Transitional Flight

The aircraft is brought to a hover and systems are checked similar tohelicopter procedures. Once final checks are complete and cleared, theaircraft is accelerated to a transition speed. At this speed, the Middlewing Ailerons, Rudder, Front and Rear Wing flaps are fully effective foraircraft control. A conversion switch is placed in the first forwardposition causing the propulsion units to move to an intermediate forwardposition. The aircraft will accelerate to a specified forward speed whenthe conversion switch is then placed in the second intermediate forwardposition which causes the propulsion units to rotate to the secondintermediate position and the wing panel flaps providing the yaw controlduring hover to return to the neutral position. Once the aircraftaccelerates to a new specified speed, the conversion switch is placed inthe forward flight position, the propulsion units rotate to the forwardflight position and the conversion process is complete. The propulsionunits' angular rate from hover to forward flight position should beapproximately 6 degrees per second.

Control Following the Loss of Propulsion Unit Thrust in Vertical Flight

When a propulsion unit failure occurs, thrust is lost causing a rolland/or pitching moment toward the failed unit. The pilot instinctivelywill move the cyclic control to correct this motion. This controlmovement increases thrust in the neighboring propulsion units anddecreases thrust in the opposite propulsion units, restoring theaircraft to its original orientation. The only thing that will change,from the pilot's perspective, is the new cyclic position in his hand.The vehicle is designed such that the center of gravity always passeswithin a polygon described by the relative positions of the remainingthrust producing propulsion units.

Control Following Loss Of Propulsion Unit Thrust in Forward Flight

When a propulsion unit fails, during departure and following transitionto forward flight, the pilot simply trims the rudder for the relativelysmall asymmetrical yaw caused by the failed propulsion unit. This yaw isnot like an engine failure in a twin engine airplane, where you losehalf your thrust and gain a large addition drag due to its wind-millingpropeller. With this configuration you lose 16.6% of your thrust andgain the additional drag of the propeller. The critically importantconcept of minimum control speed (VMC) does not exist here. The vehiclewill have accelerated beyond this speed before transition to forwardflight begins. Therefore the size, weight and cost of a large verticalstabilizer and rudder normally required for multiengine aircraft are notrequired with this invention.

Additional Features of the Invention

Propeller Blade Twist Control

There is a problem with prior art propeller technology as applied toaircraft that are designed for both static thrust (i.e. hovering flight)and high speed forward flight. More specifically, prior art propellertechnology does not effectively support such aircraft.

Traditional aircraft propellers contain blade angle distributionsoptimized for one forward speed. Accordingly, each blade station (i.e.,each point along the length of a propeller blade) is fixed at acorresponding pitch angle to achieve a best lift to drag ratio for thatone forward speed. The blade angle, though fixed at any given stationalong the length of the blade, progressively changes from a firstpredetermined blade angle adjacent the hub to a second predeterminedblade angle at the distal tip of the blade. The problem is that thisblade angle distribution typically involves the blade root chordapproaching the axis of the shaft which the propeller is mounted on.While this is optimal for the one forward speed, is not optimal forstatic thrust. If applied to a static thrust producer, much of the bladeroot would be stalled. Hence, helicopter rotor blades are sometimestwisted, but at a fraction of the twist of propeller blades.

FIG. 5 is a cross-sectional view along the longitudinal axis of apropeller 80, in accordance with an exemplary embodiment of the presentinvention, which overcomes the aforementioned problems of the prior artpropeller technology. As shown in FIG. 5, propeller 80 comprises a blade82. In accordance with a preferred embodiment, blade 82 is made ofcomposite material, such as carbon graphite, wherein the fibers arearranged such that blade 82 is flexible (i.e., twistable) about itslongitudinal axis while resisting forces parallel to the propellerspinning axis. A rod or tube 84 (e.g., a steel rod) is placed in a longhole along the longitudinal axis of blade 82, as shown. The rod 84 mayextend into the distal tip of blade 82, as shown, where it is anchored(fixed) to blade 82. In one exemplary embodiment, the distal tip of rod84 is bonded to blade 84 at or near the distal tip of blade 84. The rod84 may extend in the opposite direction into the propeller hub (notshown), where a cam 86 is attached to rod 84. A first actuator (notshown) may drive a first cam 86 which would apply a torsional force torod 84, thus causing rod 84 to rotate. It will be understood that atleast the distal tip of blade 82 would also rotate with rod 84 as thetwo are fixed to each other as explained.

Turning our attention back to blade 82, the root of blade 82 may becircular in cross section and comprise one or more blade retentionstructures 88. In accordance with preferred embodiment, the root ofblade 82, including the blade retention structures 88 are, at least inpart, mounted within an adjustable pitch blade retention mechanism. Theadjustable pitch change mechanism would include a second actuator (notshown), such as a small electric motor mounted coaxially with thepropeller. The second actuator may drive a second cam (not shown) that,in turn, may impart a force on a pin 90, positioned at the root of blade82, as shown. The force acting on pin 90 would cause the root of blade82 to twist around the rod 84. Yet, the angle associated with the distaltip of blade 82 angle may remain unchanged, as it is restrained by thefact that it is fixed to the rod 84.

The blade twist technique is similar to the technique associated withpropellers produced by IVO props. In the IVO prop design, the blade rootis fixed to the propeller hub and the torsion tube or rod is anchored tothe blade tip. The twisting of the blade is accomplished by rotating thetorsion tube or rod within the hub using a small cam, driven by aelectric motor driven leadscrew. However, in the IVO design, thepropeller blade angle at the hub is fixed and not adjustable, as in thepresent design. While the IVO design is not optimal for VTOLapplications, the present design is optimal, because in the presentdesign, the pitch angle associated with the distal tip of blade 82 isadjustable independent of the angle associated with the root of blade82, and the root of blade 82 is adjustable independent of the angleassociated with the distal tip of blade 82. Accordingly, the pitch angledistribution of blade 82 can be adjusted and, therefore, optimized toachieve effective static thrust and high speed forward flight.

Engine Size

The engine size of any version of this aircraft is significantly smallerthan competitive configurations. Smaller engines tend to be simpler tohandle and maintain. The design of most aircraft is limited to existingengine designs. It is rare that a new power plant is designed andmanufactured for a new aircraft. An aircraft which utilizes relativelysmaller engines has a much larger quantity to select from (smallerengines types are more numerous than high horse power engines). Smallerengines can be less expensive per horsepower than larger engines. Thehigher quantity of six engines allows more economies of scale to berealized. A manufacturer is far more interested if the aircraft utilizessix engines as opposed to one or two per vehicle. Quantity discounts arefar more likely in this situation.

Safety Characteristics

Historically, when an aircraft suffers a failed engine, bad things canand do happen. Single engine aircraft lose an engine and land,sometimes, successfully, sometimes not. Multiple engine aircraft maycontinue the flight, sometimes not. Engine failure on departure with atwin engine fixed wing aircraft is notoriously dangerous. Single enginehelicopters with a failed engine do not normally set down without damageor death following an autorotation. The pilot must be fast, proficient,and lucky.

The embodiment of this invention places six propulsion units inapproximately equal angular distributions around the aircraft. Thisunique placement provides for the continued flight following a singlepropulsion unit failure. The aircraft center of gravity is designed tobe located along the center line of the fuselage and approximately nearthe center of the middle wing spar. The propulsion units are designed tohave reserve power in the event of an engine propulsion unit failure.So, inherent safety following a propulsion unit failure without anyspecial pilot talent or proficiency is a unique characteristic of thisinvention.

Locating the fuel tanks within the propulsion units places the fuel awayfrom the cabin occupants. Thus, fire in the cabin during flight or crashis unlikely.

Placing the propulsion units away from the cabin prevents them fromentering the cabin during a normal crash scenario.

The forward wing stall behavior reduces the likelihood of the middle orrear wing stalling, reducing stall related accidents.

Locating the main wheels at the aft end of the forward propulsion unitswidens the space between the traditional main gear, increasing groundlateral stability.

Individual propulsion unit fuel systems provide completely independentfuel systems, preventing failure of more than one unit due to a singleproblem. There are no fuel selectors or cross-feed controls planned,just fuel shutoff switches in the cockpit.

With the advent of Primary Flight Displays (PFDs) and MultifunctionDisplays (MFD) the availability of a highly reliable power source on theaircraft is mandatory. As a matter of fact the FAA demands it. Today theFAA considers alternators and batteries as power sources. Thisembodiment shown with six propulsion units has twelve (12) powersources. A properly designed electrical system can produce the fail safepower source for these and other important electronic systems includingremote control and communication links for the UAV versions of thisvehicle. The safety increase of having a virtually fail-proof powersource is simply put, great.

Efficiencies

The rear wing of this configuration supports its share of the vehicleweight as opposed to conventional aircraft. Conventional aircraftnormally use a less efficient un-cambered airfoil known as thehorizontal stabilizer. The stabilizer operates with a downward forceduring normal flight. This force effectively adds additional weightreducing payload capability. It also creates drag due to this negativelift, known as “trim drag”. The V22 Osprey uses a conventionalhorizontal stabilizer and elevator in its configuration. The supersonicConcord actually pumped fuel to a tank in its tail to reduce this trimdrag.

The front wing is a cambered airfoil wing which operates as a liftproducer and at the same time provides pitch control as a canard withelevator. So, rather than paying for the weight, trim drag, and negativelift of the horizontal stabilizer. The rear wing becomes part of thelift producing wing function as well as sharing pitch control with thefront wing.

The location of the propulsion units near the wing ends can provide anend plate effect which reduces the normal magnitude of wing tip vorticesand improving span efficiency factor, reducing induced drag and,therefore, increasing range.

A structural efficiency can be gained by distributing propulsion unitweight near the wing tips which reduces wing root bending loads and,therefore, wing weight.

The use of six propulsion units avoids the weight, complexity, and costof a cross coupled engine drive shaft system.

Industrial Applicability

The invention is applicable to aircraft and in particular VTOL aircraft.While there have been shown and described and pointed out thefundamental novel features of the invention as applied to the preferredembodiments, it will be understood that the foregoing is considered asillustrative only of the principles of the invention and not intended tobe exhaustive or to limit the invention to the precise forms disclosed.Obvious modifications or variations are possible in light of the aboveteachings. The embodiments discussed were chosen and described toprovide the best illustration of the principles of the invention and itspractical application to enable one of ordinary skill in the art toutilize the invention in various embodiments and with variousmodifications as are suited to the particular use contemplated All suchmodifications and variations are within the scope of the invention asdetermined by the appended claims when interpreted in accordance withthe breadth to which they are entitled.

What is claimed is:
 1. A vertical takeoff and landing aircraftcomprising: a fuselage; three wings attached to the fuselage; sixsynchronously rotatable, thrust producing propulsion units located atapproximately equal angular positions about a yaw axis of the aircraft;wherein one propulsion unit is mounted above, below, or on each half ofthe three wings, and each propulsion unit is independent from the otherpropulsion units, and wherein each rotatable propulsion unit comprises apropeller and a propeller hub, and wherein each propeller comprises: aplurality of blades, wherein each of the plurality of blades comprises ahole along a longitudinal axis of each blade; and a plurality ofrotatable rods, wherein each of the plurality of rods extends into thehole of a corresponding one of the plurality of blades, wherein aproximal end of each of the plurality of blades is enclosed in oradjacent to the propeller hub and rotatable around the corresponding oneof the plurality of rotatable rods, and wherein a distal end of each ofthe plurality of rods is fixed to a distal end of the corresponding oneof the plurality of blades.
 2. The vertical takeoff and landing aircraftof claim 1, wherein each of the propulsion units further comprises: afirst actuator in mechanical communication with each of the plurality ofrods, the first actuator configured to impart a rotational force on eachof the plurality of rods; and a second actuator in mechanicalcommunication with each of the plurality of blades, the second actuatorconfigured to impart a rotational force on the proximal end of each ofthe plurality of blades.
 3. The vertical takeoff and landing aircraft ofclaim 2, wherein the first and second actuators are operationallyindependent of each other.
 4. The vertical takeoff and landing aircraftof claim 3, wherein each of the rotatable propulsion units has asubstantially vertical orientation for vertical flight and asubstantially horizontal orientation for forward flight, and wherein thefirst and second actuators are configured to adjust each of theplurality of blades so that the blades exhibit a first pitch angledistribution optimized for vertical flight and a second pitch angledistribution optimized for forward flight.
 5. The vertical takeoff andlanding aircraft of claim 1, wherein the three wings include a frontwing, a middle wing, and a rear wing, relative to the fuselage.
 6. Thevertical takeoff and landing aircraft of claim 1, wherein the proximalend of each of the plurality of blades comprises blade retentionstructures.
 7. The vertical takeoff and landing aircraft of claim 1,wherein each of the plurality of blades comprises carbon fiber compositematerial.